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Equivalent Delta-V Per Orbit of Gravitational Perturbations∗

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Equivalent Delta-V Per Orbit of Gravitational Perturbations∗ Equivalent delta-v per orbit of gravitational perturbations∗ Martin Laray Polytechnic University of Madrid – UPM, 28040 Madrid, Spain Accepted: January 5, 2016 Nomenclature a = semi-major axis (orbital element); length unit C= subindex for a term derived from the Coriolis effect c= subindex for a term derived from the centrifugal forces D= subindex for disturbing function e = eccentricity (orbital element); dimensionless f = true anomaly; rad G = angular momentum vector per unit of mass; (length unit)2=time unit I = inclination (orbital element); rad J = impulse per unit of mass (delta-v); length unit/time unit J2 = zonal harmonic coefficient of the second degree; dimensionless K= subindex for Keplerian k = unit vector in the z direction; dimensionless M = mean anomaly; rad N = reference system’s rotation rate; rad/time unit N = modulus of N; rad/time unit n = orbit mean motion n = pµ/a3; rad/time unit p = conic parameter p = aη2; length unit ∗Journal of Guidance, Control, and Dynamics, in press, http://arc.aiaa.org/loi/jgcd Copyright c 2016 by the American Institute of Aeronautics and Astronautics, Inc. All rights re- served. yResearcher. Space Dynamics Group, ETSI Aeronauticos,´ Pz. Cardenal Cisneros 3, Associate Fellow AIAA. 1 q = auxiliary symbol of order 1; dimensionless r = radius from the earth’s center of mass; length unit r = distance from the earth’s center of mass; length unit S = scalar part of the impulse in the rotating frame; length unit/time unit t = time V = gravity potential; (length unit)2=(time unit)2 x; y; z = Cartesian coordinates of r; length unit α = earth’s equatorial radius; length unit ∆v = delta-v (impulse per unit of mass); length unit/time unit = auxiliary symbol of the order of the perturbation; dimensionless p η = eccentricity function η = 1 − e2; dimensionless θ = argument of latitude θ = f + !; rad µ = earth’s gravitational parameter; (length unit)3=(time unit)2 ρ, σ = scalar indices representing the perturbed dynamics; dimensionless τ = time of perigee passage (orbital element); time unit Ω= right ascension of the ascending node (orbital element); rad ! = argument of the perigee (orbital element); rad ∗ = superindex. The indexed magnitude is valid under the averaging assumption 1 Introduction It is demonstrated that the impulse in a single orbit of an earth’s satellite can be computed analytically within a reasonable accuracy, in this way providing a vec- torial expression that is used to reveal the more important features of the earth’s satellite dynamics without need of resorting to Fourier series expansions or pertur- bation theory. The impulse is referred to a rotating frame, and is computed in the “averaging assumption”, that is, taking the osculating semi-major axis, eccentric- ity, and inclination constant in the short time interval in which the argument of the latitude of the satellite advances by 2π. In particular, by analyzing the impulse due to the perturbation that arises from the J2 term of the earth’s gravitational potential it is shown the existence of planar orbits, which may exist in the equatorial plane and also in the meridian planes (polar orbits), as well as orbits with fixed perigee at the critical inclination —all of them in agreement with well-known results. In addition, it is shown that the information provided by this impulse can be encapsulated in simple scalar indices. The evaluation of these kinds of indices can be used for the creation of argument of the perigee vs. inclination maps, which 2 provide valuable information on the perturbed dynamics. Furthermore, the proce- dure is not restricted either to the J2 perturbation or to conservative perturbations, and can be generally applied, although in most cases the quadratures defining the impulse should be evaluated using numerical integration techniques. The impulse is computed per unit of mass, which therefore will produce a cor- responding variation of the velocity vector instead of the linear momentum. Hence, it provides the equivalent delta-v per orbit of gravitational perturbations, which can be useful in practical engineering problems. For instance, some maneuvers are de- signed using just the Keplerian approximation [1]. Also, the initial design of con- stellations is commonly made under the linear effect of the J2 perturbation, which does not include the effects of the perigee dynamics [2]. In both cases a better es- timation of the delta-v budget necessary to compensate the neglected effects may be obtained from the theory provided here. 2 ∆v on a Keplerian orbit The net delta-v from Keplerian motion between two times is Z t2 µ ∆v = − 3 r dt; (1) t1 r where the time dependence of r must be made explicit to perform the quadrature. This can be done using the orbital elements representation of r. x = cos Ω cos θ − sin Ω sin θ cos I; (2) r y = sin Ω cos θ + cos Ω sin θ cos I; (3) r z = sin θ sin I; (4) r where p r = ; (5) 1 + e cos f Equations (2)–(4) and (5) reveal the time dependence of the Cartesian coordi- nates of the Kepler problem, but this is done in an implicit way through the true anomaly. Although the true anomaly is an implicit function of time, the quadrature in Eq. (1) can be solved directly either in f or θ making use of the preservation of the total angular momentum of the Keplerian motion, which can be stated in the usual form dθ p r2 = kGk = µ p = na2η: (6) dt 3 where dt = (1=n) dM, and hence r2 n dt = dM = dθ; (7) a2η making in this way available an explicit relation between the differentials of the mean anomaly and the argument of the latitude. An analytical expression of the ∆v of Keplerian motion between two arbitrary times is obtained substituting Eq. (7) into Eq. (1) to give na Z θ(t2) r ∆v = − dθ; (8) η θ(t1) r which is integrated replacing r=r by its components in Eqs. (2)–(4), namely 0 1 θ(t2) − sin θ cos Ω − cos I cos θ sin Ω na ∆v = B − sin θ sin Ω + cos I cos θ cos Ω C : (9) η @ A sin I cos θ θ(t1) As expected, Eq. (9) vanishes when it is evaluated along a full orbit θ(t2)−θ(t1) = 2π. 3 Effects of J2 −3 For the earth, it happens that J2 = O(10 ) whereas all other coefficients are of order 10−6. Then, the non-central potential ! µ µ α2 1 z2 V = − + J −1 + 3 ; (10) r r r2 2 2 r2 is quite representative of a wide class of non-resonant low earth orbits. The Newtonian equations of motion derived from Eq. (10) are written in the vectorial form " ! # d2r µ µ α2 5 z2 1 = − r + 3 J − r − zk ; (11) dt2 r3 r3 r2 2 2 r2 2 whose solutions are no longer ellipses. However, as far as J2 is small, solutions to Eq. (11) may be viewed as slightly distorted ellipses whose deformation evolves with time, a case in which the concept of osculating elements can be used in the description of the non-Keplerian motion. Thus, the position and velocity of the body at a given moment in time can be used to define an instantaneous ellipse that 4 is tangent (osculating) to the actual trajectory. Since each instantaneous ellipse is defined by corresponding Keplerian elements, the non-Keplerian trajectory can be described by the time evolution of these osculating elements, to whom all the relations of the elliptic motion apply. The osculating elements representation is quite useful to manifest the time de- pendency of Eq. (11), in this way enabling the computation of the corresponding ∆v under the averaging assumption. That is, the osculating elements a(t), e(t), I(t), Ω(t), !(t), and τ(t) are assumed to evolve slowly when compared to the fast evolution of either the mean anomaly, which is driven by the osculating mean mo- tion n(t). Since the rate of variation of the argument of the latitude will only differ from n in effects of the order of the perturbation (J2 in the present case), which are related to the non-vanishing motion of argument of the perigee, these slow varying elements are taken as constants in the computation of the quadratures defining the ∆v in the interval in which θ increases by 2π, that is, a full orbit. 3.1 The rotating frame Note, however, that while taking a, e, and I constant along a full orbit may be a reasonable approximation, the assumption that ! and Ω remain constant in the same short time interval must be further qualified. In particular, as far as the right ascension of the ascending node may have a slightly different value in the initial and final times, independently of the smallness of this a priori unknown variation, taking Ω constant is equivalent to say that the quadratures are solved in a rotating frame which is moving with the rotation rate of Ω, a case in which corresponding inertia terms should be added to obtain the ∆v. The acceleration in the rotating frame is " ! # µ µ α2 5 z2 1 r¨ = − r + 3 J − r − zk − 2N × r_ − N × (N × r); (12) r3 r3 r2 2 2 r2 2 where N = Nk is the rotation rate of the node, which is further assumed to remain constant in the averaging assumption, and overdots are used to emphasize that corresponding derivatives are measured in the rotating frame. Then, the ∆v in the rotating frame is written ∆v = J K + J D + J C + J c; (13) where the Keplerian part J K is formally the same as in Eq.
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